This invention relates generally to aircraft control systems. The invention, in particular, relates to control systems wherein the actuators of such systems, are hydraulically driven and wherein command signals from the aircraft cockpit to the control surface actuators are transmitted electrically.
For safety reasons, an attempt is made to design aircraft systems with as much redundancy as possible. An effort is generally made to provide redundancy in those areas where a failure would pose a substantial risk to the aircraft and its occupants.
Transmittal of control signals between the aircraft cockpit and the actuator for the aircraft control surface may take many forms. In small light aircraft, the signals from the aircraft pilot are mechanically transmitted to the aircraft's aerodynamic control surfaces such as flaps, elevators, rudder and ailerons. In more sophisticated aircraft the signals are transmitted hydraulically. In some cases, a hydraulic "boost" is provided so that the pilot need not supply all, if any, of the power required to move the applicable control surface.
Certain advanced aircraft have "fly by wire" control systems wherein the signals from the cockpit are transmitted electrically, to hydraulic actuator systems which mechanically move the aircraft's control surfaces.
In such systems, movement of the pilot operable member, e.g. control stick, flap handle, etc., causes movement of a transducer which generates a signal which is proportional to the magnitude of movement of the applicable pilot operable member. Such systems, depending upon their complexity, may also include force transducers, as well as other types of means for generating electrical signals. Redundancy may be provided by having two or more transducers for the same function and by providing multiple parallel electrical circuits for transmitting the signals to the control surface actuator systems.
Aircraft are generally provided with hydraulic or electrical actuators which move the aircraft's control surfaces.
It has also been found desirable to provide redundant actuators for the aircraft's control surfaces so that a failure of an actuator will not result in an inability to move the applicable control surface. Various problems arise when dual actuator redundancy is desired. Two actuators operating on a common member may work against one another if they operate at slightly different speeds or forces. These problems have been solved in some dual actuator systems by having them operate through a common mixing box, or some other such arrangement, so that such differences are mechanically compensated for in the system. These systems generally result in there being, at some point, a single load carrying member between the actuators and the aerodynamic control surface. This single load carrying pathway results in a point where there is no redundancy. This will be the weakest point in the system as the control surface itself, and the hinges associated with it, may also be provided with redundancy. In those applications where dual actuators are connected directly to the control surface the aforesaid problem is not encountered.
In some dual actuator applications, a phenomenon known as "force fight" is encountered. This occurs when the control system for the actuators themselves include position or some other parameter indicative of actuator performance, in a feedback loop in the actuator's control system. In this situation, an actuator which has arrived at its proper position may be moved by the second actuator when it is attempting to arrive at its position. This causes the first actuator to attempt to return to its proper position thereby disturbing the second actuator. This results in the two actuators operating in opposition to one another. It is because of these problems the dual rotary actuators have not been connected at opposing ends of a common control surface, particularly in "fly by wire" systems.